1. Field
The present disclosure generally relates to thermal protection materials and, in particular, to thermal protection systems using a pyrolizing ablative material.
2. Description of the Related Art
Every reentry spacecraft, such as an unmanned capsule or the Space Shuttle, requires some form of thermal protection if it is to survive the intense heat generated during atmospheric entry. Orbital or interplanetary spacecraft typically enter the atmosphere at hypersonic speed, creating a high-pressure bow shock wave in front of the vehicle. Within this shock wave, the compressed gas can reach temperatures of more than 6,000 kelvin (K), causing it to become both ionized and dissociated. This hot gas then transfers its heat to the front of the spacecraft.
One way to dissipate this large amount of thermal energy is with a heat shield on the front of the spacecraft that works by ablation, which is the process wherein the material of the heat shield melts, sublimes, and/or chemically decomposes. Phase change and endothermic chemical reactions of the heat shield material absorb energy from the shock layer gases, preventing it from reaching the spacecraft structure. Chemical reactions of the surface and near subsurface heat shield material (charring) may also produce gaseous products known as pyrolysis gases. Gas produced by pyrolysis of the heat shield material carries heat away from the spacecraft and may also block convective and radiative heat flux from reaching the spacecraft. This technique was used on the Mercury, Gemini, and Apollo programs to protect the spacecraft during re-entry in a very mass-efficient manner. Ablative heat shields provide good protection against high heat flux by a combination of chemical reactions, material phase change, and surface blowing (lifting hot shock layer gases away from the heat shield's outer wall as gases are generated at the surface of the heat shield).
One lightweight ablative heat shield material is SLA-561V (“SLA” stands for Super Lightweight Ablator), a proprietary ablative made by Lockheed Martin that has been used as the primary heat shield material on many of the vehicles sent by the National Aeronautics and Space Administration (NASA) to Mars. SLA-561V begins to exhibit ablation at a heat flux of approximately 110 W/cm2. This material is typical of most existing ablators in that it is applied to a rigid & continuous sub-structure.
The Space Shuttle orbiter uses a variety of thermal protection materials on various parts of the body. FIG. 1 shows some of the areas of the orbiter 10 where different types of material are used. Reinforced carbon-carbon 20 is used in the nose cap and leading edges of the wing where the reentry temperature exceeds 1260 degrees Celsius (C) (˜2300 degrees Fahrenheit (F)). Flexible felt Reusable Surface Insulation (FRSI) 24, bonded to the rigid underlying sub-structure, is used in areas where temperatures stay below 370 degrees C. (˜700 degrees F.).
Some of the low-temperature rigid tiles originally used on section of the fuselage, tail, and upper wing on orbiters Discovery, Atlantis and Columbia were replaced with Advanced Flexible Reusable Surface Insulation (AFRSI) in areas where the temperatures remain below 650 degrees C. (˜1200 degrees F.). AFRSI consists of a low-density fibrous batting made of high-purity silica fibers that is sandwiched between an outer woven silica high-temperature fabric and an inner woven glass lower-temperature fabric. After the composite is sewn with silica thread, it has a quilt-like appearance. Both FRSI and AFRSI are bonded to rigid sub-structure (i.e. they are not structurally self-supporting) and are limited to low heat flux levels, with FRSI generally limited to 2 W/cm2 and AFRSI generally limited to 5 W/cm2.